Gas turbine power plants are used as the primary propulsive power source for aircraft, in the forms of jet engines and turboprop engines, as auxiliary power sources for driving air compressors, hydraulic pumps, etc. on aircraft, and as stationary power supplies such as backup electrical generators or hospitals and the like. The same basic power generation principles apply for all of these types of gas turbine power plants. A gas turbine engine in its basic form includes a compressor section, a combustion section and a turbine section arranged to provide a generally axially extending flow path for the working gases. Compressed air is mixed with fuel and burned, and the expanding hot combustion gases are directed against stationary turbine guide vanes in the one or more turbine stages of the engine. The vanes turn the high velocity gas flow partially sideways to impinge at the proper angle upon turbine blades mounted on a turbine disk or wheel that is free to rotate. The force of the impinging gas causes the turbine disk to spin at high speed. The power so generated is then used to draw more air into the engine, in the case of the jet propulsion engine, and both draw more air into the engine and also supply shaft power to turn the propeller, an electric generator, or for other uses, in the cases of the other applications. The high velocity combustion gas is then passed out the aft end of the gas turbine which, in the propulsion engine applications, supplies a forward reaction force to the aircraft.
As is well known, the thermal efficiency, and therefore power, produced by any engine is a function of, among other parameters, the temperature of the working gases admitted into the turbine section. That is, all other things being equal, an increase in power from a given engine can be obtained by increasing the combustion gas temperature. This is particularly true for small turboshaft or turboprop engines where very small changes in the operating temperature can substantially affect the engine output. For example, it has been determined in a typical engine of this type that a single degree centigrade increase in the temperature of the working gases can increase the engine power by as much as 15 horsepower. However, as a practical matter, the maximum feasible gas temperature is limited by the useful operating temperature of the component parts in contact with the motive gas and/or the ability to cool these parts below the hot gas temperature.
The maximum gas temperatures occur in the combustion section. A turbine engine conventionally employs either an annular combustor or several cylindrical combustor cans arranged around the circumference of the turbine to contain the burning fuel and air and to produce energetic hot gases for introduction to the turbine section. A transition duct containing guide vanes is typically disposed between the combustors and the first turbine stage to properly direct the flow of hot gases onto the turbine wheel blades.
Various methods for cooling the walls of these combustor components have been tried in order to allow ever higher gas temperatures to be used. Most methods utilize relatively cool uncombusted air from the engine's compressor to both passively cool the exterior of the walls by convection and to actively protect the interior of the walls by film cooling.
The term film cooling as used herein refers to the technique of cooling a surface by maintaining a relatively slow moving layer or film of cool air near the surface so that the layer acts as an insulative barrier to prevent or retard unwanted heating of the surface by the adjacent hot gas stream. In this context, film cooling is distinguished from the more common convection cooling which operates on the completely different principle of maintaining a relatively high velocity flow of cooling air at a surface to carry heat away from the surface rather than insulating the surface from an adjacent heat source.
Several problems exist with the known cooling methods when applied in smaller high performance gas turbine engines. Simple film cooling through slots and/or louvers in the combustor walls does not utilize the full heat sink potential of the cooling air. Also the amount of air so used represents a significant portion of the total air flow from the compressor which would otherwise be available to support combustion and control the burner exit temperature profile, i.e. eliminate hot spots.
To use cooling air more efficiently, recent attempts have focused on providing film cooling through arrays of holes or passages, as opposed to continuous slots, and constructing the passages to provide more active internal wall cooling by convection or impingement, or both. See, for example, U.S. Pat. Nos. 3,420,058; 3,623,711; 3,737,152; 4,242,871; 4,622,821 and 4,773,593. Such complicated cooling schemes raise new problems to be solved. For example, the uniform hole patterns normally employed can result in wall sections that are undercooled on the leading (upstream) edge, well cooled in some central regions, and overcooled on the trailing edge as the cooling film effectiveness increases from row to row in the streamwise direction. In addition, the typical pressure drop through the combustor walls tends to produce blowing ratios much larger than the ideal value near 0.4. Hence, the effusion jets can separate from the hot surface and mix with the bulk flow rather than forming a protective film near the surface.
In view of the foregoing, it is an object of the present invention to provide an improved cooling system for gas turbine combustor walls.
More specifically, it is an object of the invention to provide a durable but lightweight combustor liner having a more effective array of cooling passages therein.
It is a further object of the invention to provide an efficient method of making a porous combustor liner for advanced gas turbine engines.